It is well-known that the exposed frontal area of a jet aircraft generates a rapidly increasing drag penalty as the aircraft achieves supersonic flight. Because a significant portion of the exposed frontal area is often directly attributable to the inlet size of the gas turbine engine employed to power the jet aircraft, it is apparent that the inlet should be minimized to reduce the drag penalty generated during supersonic flight. A reduced gas turbine inlet can provide sufficient air flow at supersonic speeds due to the relatively high ram pressures generated by the aircraft. However, the same gas turbine inlet cannot provide adequate air flow during take-off and low speed flight due to the greatly reduced ram pressures generated during slow speed flight of the aircraft.
In an effort to solve the problem of providing sufficient air flow through a gas turbine inlet during subsonic and supersonic flight, it has been suggested that additional air flow passageways be constructed for selective use only during take-off and low speed flight operations. Preferably, the air passageways would be formed in the inlet portion of the engine nacelle assembly, thereby allowing additional air to enter the gas turbine inlet. In order to maintain the required pressure level in the inlet during supersonic flight, it would be necessary to close the air passageways to prevent air pressure leakage therethrough. At this point, it was suggested that door assemblies be mounted in the nacelle for selectively covering and uncovering the air passageways as required. During take-off and low speed flight, the door assemblies would be selectively moved to open positions, allowing additional air to flow through the air passageways and enter the gas turbine inlet. As the flight speed of the supersonic jet aircraft increased to a level wherein the ram pressure provided sufficient air flow through the relatively small gas turbine inlet, the door assemblies would be selectively moved to their closed positions. This would prevent the relatively highly pressurized air from escaping from the inlet through the air passageways. However, if excessive air pressure were to be generated within the gas turbine inlet, the door assemblies could be selectively moved to their open positions, thereby allowing the excessive air pressure to be vented from the inlet via the air passageways.
While such a system would appear to provide a means for controlling air flow pressure in a supersonic gas turbine inlet, known systems have proven generally unsatisfactory in preventing excessive propulsion system losses, especially during supersonic flight operations of the aircraft. Even though conventional door assemblies attempt to block air flow through the air passageways leading from the gas turbine inlet, a stream of air tends to leak past known door assemblies, decreasing the operating pressure in the inlet and significantly reducing the operating efficiency of the power plant during supersonic flight.
As will become clear, the present invention provides an improved system for controlling air flow through a gas turbine inlet during both subsonic and supersonic flight, while at the same time providing a unique, compression-sealed door assembly for preventing excessive leakage of inlet air from the gas turbine even during supersonic flight operations of the jet aircraft.